48th International Conference on Environmental Systems ICES-2018-148 8-12 July 2018, Albuquerque, New Mexico
Sentinel-2A/B Thermal Design - Lessons Learnt from TBTV, LEOP and IOC
Nadine Buhl1, Martin Altenburg2, Markus Manns3 Airbus Defence and Space GmbH, Claude-Dornier-Straße, 88090 Immenstaad , Germany
The Sentinel-2 (S-2) satellite is developed by ESA and forms part of the EU Copernicus program. It carries an innovative wide swath high-resolution Multispectral Imager for a new perspective of our land and vegetation. The combination of high resolution, novel spectral capabilities, a swath width of 290km and frequent revisit times will provide unprecedented views of Earth. The mission is based on a constellation of two identical satellites in the same orbit, 180° apart for optimal coverage and data delivery. Together they cover all Earth’s land surfaces, large islands, inland and coastal waters every five days at the equator. S-2A was launched June 2015 and S-2B March 2017. The thermal design has been analyzed and verified through thermal balance and vacuum tests to validate the thermal mathematical model. The paper describes the development of the mathematical model, the execution of the thermal tests in different facilities (IABG and ESTEC) and the corresponding correlation activities. The predicted and actual in-orbit performances are also reported. The thermal balance performed with the S-2A satellite showed already a good match of the thermal model to the as-built status. Nevertheless, some discrepancies were found on the Star Tracker Assembly and the Propulsion Plate. The thermal test of the S-2B satellite revealed further needs to improve the S-2A correlated thermal model. A summary of the most relevant parameters is given within this paper. The comparison of flight data during the in-orbit commissioning phase with the flight predictions performed with the correlated model show discrepancies in the order of only +/-3K for most units. The comparison of duty cycles shows deviations higher than 10% only for the units which already had higher discrepancies during the thermal balance correlation, namely the Star Tracker Assembly and the Propulsion Plate. The comparison of flight data shows a good similarity in temperature, stability and performance of both satellites. Continuous monitoring of the thermal behavior since launch is also revealing noticeable degradation of thermo-optical properties due to sun exposure, in particular on the Star Trackers.
1 Thermal Engineering & AIT, Thermal Engineering FHN, [email protected] 2 Thermal Analysis and Design Expert, Thermal Engineering FHN, [email protected] 3 System Architecture & IV&V, Mechanical Systems Germany, [email protected]
Copyright © 2018 Airbus Defence and Space GmbH
MMFU = Mass Memory and Formatting Unit Nomenclature MSI = Multi Spectral Instrument BOL = Beginning Of Life OBC = On-Board Computer CDR = Critical Design Review OCP = Optical Communication Payload CESS = Coarse Sun and Earth Sensor PCDU = Power Control and Distribution Unit CFRP = Carbon Fiber Reinforced Plastic PDR = Preliminary Design Review EOL = End Of Life PFM = Proto Flight Model ESA = European Space Agency PROP = Propulsion ESTEC = European Space Research and Technology RIU = Remote Interface Unit Centre S-2 = Sentinel-2 FM = Flight Model S/A = Solar Array FPA = Focal Plane Assembly S/C = Spacecraft GEU = Gyro Electronic Unit (IMU electronics) SADM = Solar Array Drive Mechanism GFU = Gyro-Fibre Unit (IMU sensor) SBT = GPS Satellite Binary Time GL = Linear Coupling STR = Star Tracker GMM = Geometrical Mathematical Model TB = Thermal Balance GPS = Global Positioning System TBTV = Thermal Balance and Thermal Vacuum GPSR = GPS Receiver TCS = Thermal Control System GR = Radiative Coupling TICD = Thermal Interface Control Document IABG = Industrieanlagen-Betriebsgesellschaft mbH TMM = Thermal Mathematical Model IMU = Inertial Measurement Unit TV = Thermal Vacuum IOC = In-Orbit Commissioning VCU = Video and Compression Unit IOV = In-Orbit Verification VDA = Vacuum Deposited Aluminium LCT = Laser Communication Terminal WSA = Weltraumsimulationsanlage (Space LEOP = Launch and Early Orbit Phase Simulation Chamber) LSS = Large Space Simulator XBS = X-Band Subsystem MLI = Multi Layer Insulation I. Introduction
HIS chapter provides a comprehensive overview of the S-2A/B satellite design with focus on the platform side. T The satellite is characterized by a modular configuration enabling as much as possible parallel development and integration of its main subassemblies. The main structure of the platform is constituted by an aluminum frame structure with an intermediate floor, which divides the satellite in a lower electronics and an upper instrument compartment. The instrument compartment is closed by 4 access/radiator panels on Y&Z-sides. The radiator areas are optimized to the minimum possible sizes, reducing the need for operational heater power in cold case conditions. All external surfaces not used as radiators are thermally insulated to the maximum extent possible by means of Multi-Layer Insulation (MLI) blankets, limited by a number of unavoidable heat leaks, as there are the launcher I/Fs, thrusters, sensor apertures and antennas. An electrical heater system consisting of independently controllable heater lines is implemented to maintain minimum unit and payload temperatures during operating and non-operating phases as well as for temperature stabilization of the Star Tracker (STR) and Giro Fibre Unit (GFU) assemblies.
Figure I-1: Satellite Design Overview.
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The top floor, on +X-side, is providing accommodation for the Multi-Spectral Instrument (MSI), as shown in Figure I-1, which is iso-statically mounted by means of six CFRP struts on top of an intermediate CFRP panel which is again attached by 9 brackets to the platform top panel. The MSI electronic units are mounted close to the instrument on the rear side of the top panel. To achieve the required pointing accuracy, the alignment critical sensor assemblies carrying Star Trackers and IMU sensor (GFU) are accommodated directly on the MSI, as shown in Figure I-2. The Platform unit equipment compartment accommodates (by means of 4 sandwich panels and the intermediate floor) the electronic equipment. The satellite units are mounted on the inside of the platform aluminium sandwich panels with exception of the batteries which are mounted thermally insulated on the outside allowing for independent thermal control and late integration. The units have been distributed to equalize as much as possible the thermal dissipation, but also to keep the center of gravity balancing required by the launcher. The main dissipating units are the VCUs, PCDU, RIU and MMFU. The X-Band Subsystem as a whole also exhibits a substantial orbit average as well as peak heat dissipation.
Radiators on Zenith and Deep Space side Radiators on Nadir and Sun side
STR Radiators SBT & GEU Radiator VCU Radiator Access Panel Radiator
Access Panel Radiator SADM Radiator LCT Radiator Battery Module Radiators
OBC Radiator MMFU Radiator
PCDU Radiator XBS Radiator GPSR Radiator WDE Radiator
Figure I-2: Satellite External Thermal Model. . A milled aluminium base plate is providing the interface to the cylindrical adapter with the clampband I/F connecting the satellite to the launch vehicle adapter. The propulsion subsystem is designed as a separate module which will be fully integrated prior to the integration into the platform. It is thermally decoupled from the satellite structure to the maximum extent possible. Since the minimum operational temperature of the RCS equipment is higher than the minimum operational temperature of the other internal units the propulsion module is covered with MLI to reduce radiative heat exchange and to minimize heat power for cold case. The pipework lies within an enclosure formed by the propulsion plate and the external MLI. Its temperature is radiatively controlled by means of heaters installed on the propulsion plate. The single wing solar array is designed such that its rotation axis is in line with both center of gravity of the S/C and the deployed wing, which minimizes the in-flight disturbances. It is considerably reducing the heat rejection capability of the PCDU and SBT/GEU radiators.
II. TB/TV Test
The Sentinel-2 satellites were subjected to thermal balance (TB) and thermal vacuum (TV) tests. The TB test is only performed on the S-2A PFM satellite as part of a combined TB/TV test while the S-2B FM satellite is just subjected to a TV test. However, in order to verify the performance and workmanship of the TCS, at least one FM TV test
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phase had to be executed in exactly the same way as on the PFM. The two test campaigns were performed at IABG and ESTEC respectively. The S-2 satellite level thermal test objectives are summarized below:
Thermal Balance Test Objectives: qualification of the platform thermal design & hardware w.r.t. performance and workmanship validation of the Thermal-Mathematical Model (TMM) of the Sentinel-2 platform by means of steady-state (thermal balance) as well as transient (orbit simulation) test phases verification of platform to payload / unit thermal interfaces confirmation of radiator and heater sizing already verified on instrument/equipment level
Thermal Vacuum Test Objectives: verification of the functional performance of the satellite under extreme in-orbit environmental conditions by means of Abbreviated Function Tests and Special Tests at satellite qualification (S-2A PFM) or acceptance (S-2B FM) temperature levels verification of the function and performance of the on-board thermal control S/W verification of the satellite workmanship
The S-2A TB/TV test profile in comparison to the S-2B TV profile is shown in Figure II-1. The adjusted S/C temperature level, chamber shroud temperature, sun simulator intensity and chamber pressure are presented. In addition, it is indicated in which test phases functional tests are performed. The unit and structure TV target temperatures are in general derived from the flight predictions (incl. uncertainty margin) by adding ±10K qualification or ±5°C acceptance margin respectively. During cold phases the unit temperatures were allowed to go down to the minimum acceptance and during hot phases up to maximum acceptance limits. However, due to test and design limitations it was not possible to reach all unit target temperatures without the risk to stress or damage other components.
S-2A PFM S-2B FM2
0
no sun simulation
Figure II-1: S-2 PFM and FM2 test profile.
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The different chamber configurations for S-2A and S-2B are shown in Figure II-2. In the Weltraumsimulationsanlage (WSA) at IABG, for test phases without sun, the shutter remains closed. Within the Large Space Simulator (LSS), the S/C has at every test phase a view factor to the mirror, since no shutter is available. The WSA model includes the L-adapter, on which the S/C is attached to via the thermal test adapter. This adapter includes an embedded thermal shroud to set and regulate needed boundary temperatures for the propulsion module. In the LSS, the thermal test adapter, including the embedded shroud, is mounted on top of the gimbal stand (shroud 3). In both test configurations, test equipment has been modelled as far as relevant for model correlation, like the GPS re-radiation antenna (S-2A TB/TV test only), the SADM test harness support bracket and power harness.
WSA chamber with closed shutter LSS
Figure II-2: Sentinel-2 in the Vacuum Chamber. Left S-2A at IABG, Right S-2B at ESTEC
The test equipment temperatures measured during the TB phases are set as boundary conditions for the S-2A thermal model correlation. This means that the test equipment modelling has not been modified during the correlation, and possible deviations of predicted temperatures before the TB/TV test with respect to measured values are neglected. This is applicable since the test equipment is non-flight hardware, hence has been removed from the S/C completely after the TB/TV test. The equilibrium criterion for the 2 TB phases was defined as a maximum temperature drift of 0.5 K over a time period of 5 hours, whereas the stabilization criterion for TV phases was relaxed to a maximum temperature variation of 1K over 2 hours.
The S-2A test correlation runs were computed with the S-2A GMM and TMM in ESATAN TMS r6 source code with begin of life (BOL) optical properties, which was modified to respect the specific changes that are necessary for the test set-up. The main modifications of the spacecraft part according to the test configuration are listed below: Removal of the Solar Array Introduction of the test MLI (-Y radiator and +Z radiator panel) Modification of the set points of the platform heaters Introduction of actual unit dissipation values during test Introduction of the Test Heaters Introduction of the Ground Handling Adapter Introduction of the Thermal Test Adapter with embedded shroud
The overall interface couplings, i.e. couplings between all unit side panels and the frame, as well as all other frame couplings (to baseplate, to Intermediate floor, etc.) have been increased by 20%. Additionally, the overall MLI efficiency of all blankets (except SLI and MSI MLI, since already correlated), has been decreased by 30%. This means that the conductive and radiative couplings through the MLI layers have been increased. The in-plane conductivity of all unit side panels and access panels, as well as both radiator panels has been increased by 30%. This is justified by the fact that inserts have not been taken into account in the original modelling. Further, local increases of the thickness as well as stiffeners had not been considered so far.
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The local effect in heat conductivity and heat distribution due to the sum of those effects is now represented through the increased in-.plane conductivity, leading to a more proper temperature distribution especially on adjacent nodes. The unit temperature reference points are not affected by that modification. Finally, the PCDU contact conductance, initially assumed to be 300 W/m² has been increased by 20% up to 360 W/m².
Major temperature deviations at the STR assembly and GFU assembly were observed during Hot TB compared to the test prediction performed previously. Hence, some modifications of the STR assembly and GFU assembly modelling had to be performed. The MLI modelling has been modified to consider reduced MLI efficiencies, gaps between structure and MLI tent or the local heat path through the stand-offs. To reduce local deviations, for example on the STR radiators, correlation factors have been implemented. The correlation factors are in a range between 0.5 and 2.0.
In the original modelling, the Coarse Earth and Sun Sensor (CESS) was not explicitly modelled, since its effect on the GFU assembly was considered negligible. During the correlation it was found that the major discrepancies are partially due to the missing CESS and the heat exchange associated to it. Further, the GFU has been introduced, to take into account the internal radiative environment of the GFU assembly and to simulate the heat path through the mechanical interfaces. The introduction of the GFU leads to significantly better correlated temperatures in the GFU assembly. Also some minor faults in the modelling of thermal optical properties on a small number of shells (MLI) were corrected. Conductive couplings within the GFU and STR assemblies as well as between them and the secondary structure of the MSI have been introduced to simulate the heat path through electrical grounding and harness. The following Figure II-3 gives an overview of the main grounding straps (pink colored). The influence of harness form the GFU, STRs, heaters including thermistors and also test implementation is now reflected.
Figure II-3: Main grounding straps on STR assembly and GFU assembly
Major temperature deviations at the Thruster alignment brackets were observed during Cold TB compared to the test prediction performed previously. Hence, some modifications of the propulsion module modelling had to be performed. In a first attempt, the MLI modelling has been modified. So far, the baseplate SLI had been modelled without considering possible fitting constraints due to the harness routing and connectors at the baseplate. Due to the harness connections, it is not possible to close the baseplate SLI completely. This has been implemented in the GMM by deleting the corresponding SLI surface, which leads to a view factor of the baseplate structure to the propulsion MLI.
Further, conductive couplings between the Propulsion Plate and the external layer of the MLI have been introduced, simulating the heat path through the stand-offs. Also, conductive couplings between propulsion plate and pipework have been implemented, which had been neglected so far. 6 International Conference on Environmental Systems
The overall MLI coupling of the propulsion module has been decreased by 50%, the in-plane conductivity of the Propulsion Plate has been increased by 50% to account for local reinforcements and the interface coupling of the propulsion plate to the cone has been increased by 30%. Regarding the thruster alignment brackets, the thermal strap has been remodeled slightly, leading to modifications in the conductive coupling from alignment bracket to the thruster support bracket via the strap. Additionally, the MLI efficiency has been decreased, and the conductive coupling from the inner MLI layer to the alignment bracket has been increased since there is only a small gap between both. The combined effect of all modifications explained in this section leads to a better correlating thruster alignment bracket temperature as well as overall propulsion module temperature level.
Harness routing Heater and sensor locations on the propulsion module Propulsion Module T_PT1 T_PT3 T_PROP-1/2/3 T_PT2 T_TC_COLD-1/2
H_PROP_1A/B
H_PROP_2A/B
T_PROP-4/5/6
H_PROP_4A/B
T_PROP-10/11/12
T_PROP-13/14/15
H_PROP_5A/B
Thruster A1 … B4
T_FCV-xy H_FCV-x_A/B
T_PROP-7/8/9 H_PROP_3A/B
Note: = Unit TRP (Note: TC_COLD-1/2 mounted on RCS CB) Figure II-4: Complex thermal design on the Propulsion Plate.
The predicted and measured duty cycles of the STR assembly, GFU assembly and Propulsion Plate heater lines, shown in Figure II-4, are summarized in Table II-1. Only these heater lines are considered, since during the TB correlation, major model modifications have been performed. The most significant deviations are found in the H_PROP_1 and H_PROP_3 heater lines, as well as in the H_STRA_XP and H_STR-1/2/3 heater lines. Although the predicted duty cycles deviate approx. 10-30% from the measured values, the predicted ones are in most cases higher, hence, leading to worst case results.
Table II-1: Hot Orbit Simulation Heater Duty Cycles
Heater Heater Duty Cycles [%] Designation Predicted Measured H_GFA_BP 79.25 58.23 H_STA_BP 62.85 62.73 H_STA_XM 46.27 37.94 H_STA_XP 44.43 69.22 H_STA_YM 30.83 22.46 H_STA_ZM 53.84 59.07 H_STR-3 16.29 5.82 H_STR-1 18.80 3.33 H_STR-2 33.33 43.43 H_PROP_1 37.22 0.00 H_PROP_2 23.79 27.29 H_PROP_3 17.36 0.00 H_PROP_4 16.53 11.31 H_PROP_5 15.84 15.31 7 International Conference on Environmental Systems
The following update of the S-2A correlated model are performed to improve the results of the S-2B TV test correlation, which show major deviations in GFU assembly and STR assembly predicted temperatures w.r.t measured values using the S-2A correlated model. In the frame of the S-2B TV correlation, higher discrepancies between measured and predicted temperatures were found on the STR assembly and GFU assembly. Therefore, a number of model modifications were performed to improve the correlation between thermal model and real hardware. In addition to the STR assembly and GFU assembly, the Propulsion Module was also slightly modified. The modifications for each subsystem are outlined in the following sections. Only the final parameter settings are outlined; the individual steps to achieve each parameter set are not presented within this document, since they are not of importance for the final result.
During the model update of the GFU assembly, it was found out that the CESS on the GFU cylinder was not properly modelled. The CESS cover was set to inactive, so that in the Cold TB analysis case, the cylinder structure of the GFU assembly had incident solar fluxes. In addition, the cylinder top cover had direct view factors to the cold shroud. This was corrected, but after a re-run of both S-2A Hot and Cold TB analysis cases, higher temperature differences than prior to the modification were observed. Especially the Cold TB presented higher deviations, leading to significantly lower predicted temperatures than measured, with a delta of approx. 10K between them. As a consequence, the correlation performed in the frame of the S-2A thermal test had to be re-assessed completely. Reasonable modifications in thermal conductive couplings had a positive impact on the model correlation but were not enough to meet the required accuracy; therefore, the MLI modelling was re-assessed and major modifications were performed, as summarized hereafter.
The parameter modifications of the STR assembly GMM/TMM for S2-A are: Absorptivity of the MLI increased from 0.04 to 0.16 since the assumed alpha was too low Specularity of the MSI MLI outer layer (Kapton VDA) set to 0.3 specular and 0.35 diffusive (0.35 solar absorbed and 0.65 reflected) Size of STR radiator holder MLI surface increased in the GMM MLI efficiency factor of the STR assembly bracket MLI is set to 4 MLI efficiency factor of the STR to the STR assembly MLI is set to 1.5 MLI efficiency factor of the STR housing to MLI set to 2 for all 3 STRs Linear coupling of STR 2 housing to radiator reduced by 20%
The parameter modifications of the GFU assembly GMM/TMM for S2-A are: Absorptivity of the MLI increased from 0.04 to 0.26 on the top of the GFU assembly cylinder and 0.16 on all other MLI blankets since the assumed alpha was too low Specularity of the MSI MLI outer layer (Kapton VDA) set to 0.3 specular and 0.35 diffusive (0.35 solar absorbed and 0.65 reflected) MLI efficiency factor set to 1.2 MLI efficiency factor of GFU assembly to MSI MLI set to 2 Linear coupling of GFU assembly baseplate to MSI I/F factorized by 1.5
These parameter modifications led to acceptable temperatures for the analysis cases without sun simulation. The highest discrepancies between measured and predicted temperatures are observed in the Cold TB, where the sun simulator was ON.
Similarly as for STR and GFU assembly, in the frame of the S-2B TV correlation, a major temperature discrepancy between measured and predicted values was observed again for the propulsion module. The thermistor triplet PROP-4/5/6 is located on the pipework in the vicinity of the connector bracket, which is in general colder than the propulsion plate. A harness coupling of 0.002 W/K was introduced between the thermal node corresponding to the PROP4 and the connector bracket. In the frame of the S-2B TV correlation, major temperature differences between measured and predicted values were observed in the GFU and STR assemblies. As a consequence, the S-2A correlated model was updated to fit with both TB (done with S-2A) and TV (done with S-2B) cases. The test cases without sun simulator presented in general a better correlation with the test results. The highest deviations were observed in the Cold TB, where the sun simulator was active. This leads to the conclusion that the sun has a major impact, which was not reflected so far in the S-2 model. 8 International Conference on Environmental Systems
Several modifications were performed in the GMM in terms of optical property and surface re-modelling and in the TMM in terms of MLI efficiencies mainly. The presented final S-2B correlated model version provides good results for all test cases analyzed for the S-2A TB and S-2B TV test. Any further modifications in the TMM, such as MLI efficiencies or couplings between parts of the subsystems, lead to better results in the hot cases, but worse results in the cold cases. Hence, the current model status provides a good compromise between all relevant thermal test cases. A sensitivity on the test heater power was performed leading to the conclusion that a variation in applied test heater power has significant impact in the resulting temperatures. The incertitude in exact test heater power values is driven by the test facility data handling system and the observed offsets in the heater lines during phases where the heaters were off. Some of the temperature differences in the correlated model can be explained by the small inaccuracy of the test heater power.
III. Launch and Early Orbit Phase and In-Orbit Verification
S-2A was launched in June 2015 and its twin S-2B in March 2017. In terms of orbital conditions, a launch in June means a cold environment in-orbit, whereas a launch in March corresponds to a warmer environment, i.e. Earth Infrared and albedo values between hot and cold conditions. The temperature of the S/C on the launch pad was from 18°C to 20°C. Thus, this is the starting temperature from which the S/C components slowly decrease during LEOP and commissioning until achieving stable temperature conditions in-orbit, ideally similar to the flight predictions. The duration of the In-Orbit Verification (IOV) of S-2A was from Mid/End June until Mid-August, whereas the IOV duration of S-2B from beginning of March up to Mid/End of April.
According to Vega’s user manual, the S/C is exposed to a considerable thermal heat flux after jettison, due to the 3rd stage (Z9) engine firing. Several analyses have been performed beforehand and afterwards to assess the impact on the Solar Array and the CESS, located close to the launcher I/F. One concern was the effect of too high temperatures on the outer layer of the MLI enwrapping the S/C. A comparison of the in-orbit measured temperatures with the predicted temperature profiles suggested that the incident plume IR fluxes were actually significantly lower than specified. A maximum temperature of approx. 40°C was observed on the CESS facing the anti-flight direction; hence, critical temperatures have not been reached on any external MLI surface. The actual measured temperatures on the CESS facing the anti-flight direction and the timeline of the plume flux is shown in Figure III-1 for both S-2A and S-2B. A detailed evaluation of the S-2A in-orbit data is provided in Ref. 1. The findings from S-2A were generally confirmed on S-2B.
Figure III-1: Measured CESS Temperatures during Launch – S-2A (left) and S-2B (right).
After successful launch and achievement of stable power conditions in-orbit, i.e. solar array deployment, the nominal thermal mode table of the Thermal Control System (TCS) was activated, triggering the switch-on of additional platform heaters to maintain equipment temperatures above their minimum operating limits. After detailed evaluation of the S/C in-orbit thermal behavior, several heater set-points were adjusted to optimize the heater power distribution and the switching cycles.
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The modifications were performed on the propulsion system heaters, which already showed during the thermal model correlation some discrepancies with respect to the predictions. The TCS fine tuning performed in the frame of S-2A regarding heater set-points has been confirmed in general on S-2B, with some minor differences in the switching behavior of the propulsion module heaters.
During the LEOP and IOV phase, only one relevant TCS anomaly was observed for S-2A. The thermistor on the middle panel of the Solar Array (S/A) developed an open circuit failure in a specific temperature range, leading to erroneous measurements on low temperatures (below -10°C). This effect was not present at beginning of life in orbit, but appeared soon afterwards. Other signals transmitted via the same connector are nominal. Since the actual S/A temperature is not used operationally, no actions needed to be taken regarding this anomaly. In addition, the actual temperature value still could be deducted from the thermistors located on the neighboring solar array panels.
Another unexpected behavior was, that it was noticed right from the beginning that the MSI Focal Plane Assembly (FPA) temperature and required stability could not be maintained with the thermal control system. The design had foreseen to compensate fluctuations in incident fluxes by means of a heater and achieve a stable temperature around the defined set-point by heating the FPA at a certain interval. Nevertheless, the temperature regulation was lost over a significant part of the orbit. This means that the temperature variations of the FPA were higher than expected and the heater showed a duty cycle of 0%. It was considered to be caused by a considerable variation of Earth infrared flux / temperature and albedo with the latitude. This effect had not been taken into account during the thermal analysis, since the flight predictions were performed based on uniform Earth infrared and albedo distributions. This hypothesis was reinforced by the observed in-orbit temperature profile of the CESS having the same view factor than the FPA. In order to regain the full temperature regulation on the FPA, the heater set-point was increased by a couple of degrees Celsius, leading to the required temperature stability over orbit.
Regarding S-2B, an interesting Solar Array temperature anomaly occurred soon after S/A deployment. It was noticed that one of the S/A panel thermistors was not behaving as expected during the sunlit phase of the orbit as it was showing a lower temperature than the thermistors on the other two S/A panels (inner and outer one) while S/A panel 1 should be the coldest one due to its larger view factor to space. Based on this observation a delamination, or at least a poor thermal coupling of the thermistor to the S/A panel, was supposed to be the cause for this behavior as the power system was behaving nominally. A series of thermal analyses were performed to investigate this occurrence. For this, the thermistor was modelled as a specific thermal node and the thermal coupling (GL) between this node and the S/A panel was varied to find out the effect on the predicted temperature. The lower the thermal coupling, the higher the temperature offset of the thermistor. Figure III-2 shows the comparison of predicted S/A temperatures and thermistor 2 reading with in-orbit measured data. The assumed GL was 0.0035 W/mK, leading to an offset of the thermistor reading of approx. 8K w.r.t. to the panel temperature. A GL of 0.0015 W/mK lead to an offset of approx. 16K, which was already too high compared to the measured in-orbit data.
Figure III-2: Comparison of Predicted and In-Orbit Solar Array Temperatures.
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However, it was later on observed that after a modification in the bus voltage of the S/C, the S/A panel 2 temperature offset in regard to the other two thermistors on panel 1 and 3 had vanished as soon as no power was drawn from the SA. Therefore, the most probable explanation of the observed SA Panel 2 temperature offset is a non-uniform power generation of the S/A. For comparison of in-orbit observed temperatures with flight predictions, a period with sufficiently stable conditions and settled temperatures had to be selected. A dedicated analysis case was performed matching the real in-orbit conditions at that time in terms of Earth temperature and albedo, as well as S/C configuration. Figure III-3 shows some representative component temperatures for S-2A and S-2B during the IOV. For S-2A, the selected time span was from 1st to 2nd of July, and for S-2B from 12th to 13th of March. The temperature prediction (without uncertainty) is outlined as a blue bar chart and corresponds to the same thermal analysis load case for S-2A and S- 2B, even though actual environmental conditions are not identical. The correlation of in-orbit measured (orange) and predicted temperatures (blue) reveals a maximum discrepancy of 6K. However, most of the units show a discrepancy within 3K, which is well within the individual temperature uncertainty margin. This confirms the results of the Thermal Balance test correlation for S-2A and the Thermal Vacuum test correlation for S-2B. In general, warmer temperatures for S-2B are observed. This is due to the fact that the launch from S-2B was in March and S-2A in June. In March, the solar constant, the Earth infrared and the albedo is higher than in June, leading to higher incident fluxes on the S/C and consequently, higher equipment temperatures.
Figure III-3: Comparison of Predicted (blue) and Measured (orange) Equipment Temperatures in [°C] during IOV. Top: S-2A, Bottom: S-2B.
Table III-1: Comparison of Predicted and In-Orbit Measured Heater Duty Cycles in [%] during IOV. S-2A S-2B Heater Line In-Orbit Predicted Delta In-Orbit Predicted Delta BAT 23.9 25.9 -2.0 24.5 26.8 -2.2 GFA BP 44.5 60.2 -15.7 49.3 67.0 -17.1 PROP-1 3.5 36.6 -33.1 19.7 35.3 -15.6 PROP-2 55.0 31.0 24.0 49.4 39.0 10.5 PROP-3 14.9 43.6 -28.8 0.0 35.1 -35.1 SADM 15.3 6.9 8.5 6.1 7.5 -1.5 STA BP 71.5 61.4 10.1 81.0 63.5 17.5 TANK 31.9 25.0 6.9 19.4 11.9 7.5
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Regarding the heater switching, the majority of the duty cycles show differences lower than 10%. Differences greater than 10% were found for the Propulsion System, Star Tracker Assembly and Gyro Fiber Unit Assembly, deviations which were already observed during the TB correlation. In most cases, the duty cycle in-orbit is lower than predicted. This means the flight prediction is conservative. As a result of the heater duty cycle evaluation on S-2A, the set-points of the propulsion heater lines were slightly modified to achieve a more uniform heater power distribution on the propulsion plate and to reduce the temperature amplitudes.
IV. Long Term Performance Analysis and In-Orbit Comparison of S-2A and S-2B
All platform units of S-2A/B are well within their allowed temperature limits. Daily temperature variations are low for internal units as far as nominal operations are performed. Slight temperature drifts are visible due to changing environmental conditions, such that an overall temperature increase is observed during the transition from summer to winter, and vice versa during the transition from winter to summer. The following comparison of flight data reveals a good similarity in temperature, stability and performance of both satellites.
Platform temperatures are highly depending on the actual overall power dissipation of the electronic units. When comparing the platform temperatures on S-2A and S-2B, the respective total S/C power demand needs to be taken into account.The comparison period outlined in this paper corresponds to the period from September 2017 up to November 2017. During this phase, the power consumption of both S/C is nearly identical, as can be seen in Figure III-4.
Figure III-4: Comparison of Total In-Orbit S/C Power Demand between S-2A and S-2B.
The comparison of representative equipment panel temperatures of S-2A and S-2B shows very similar conditions on both S/C, pointing out the similarity of the design and build status, as outlined in Figure III-5 and Figure III-6 for average panel temperatures, in which YP refers to the panel pointing towards Sun and YM refers to the panel pointing towards Deep Space. The S-2B temperatures still show a slight offset with respect to the S-2A temperatures due to the fact that S-2B was launched only 6 months prior to the investigated period, whereas S-2A is already two years in orbit and the temperatures have reached steady-state. Nevertheless, the S-2B temperatures are approaching the levels of S-2A; at some panels, the temperature levels are already almost coincident at the end of the considered period.
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Figure III-5: Comparison of Average In-Orbit Panel Temperatures between S-2A and S-2B.
Figure III-6: Comparison of Average In-Orbit Panel Temperatures between S-2A and S-2B.
The already mentioned discrepancies on the propulsion module between the as built S/Cs and the thermal model, as well as the differences between S-2A and S-2B are clearly visible in Figure III-7. The overall temperature level is similar, but the heater switching between both S/C is different, leading to different temperature fluctuations on the propulsion plate. The observed fluctuations are in general lower on S-2B as only 2 out of 3 propulsion plate heater circuits are actually switched on and off while the third remains continuously off. On S-2A all 3 propulsion plate heater circuits are switching. The different heater behavior on both S/Cs shows that even small variations in configuration, such as a slight different thermistor positioning, can have a visible impact on the duty cycles. Even though the sum of all duty cycles is the same, the individual heater lines are different from S-2A to S-2B. 13 International Conference on Environmental Systems
Figure III-7: Comparison of Propulsion Module Temperatures between S-2A and S-2B.
One additional discrepancy between both satellites is found in the Star Tracker Assembly. Figure III-8 shows the Star Tracker temperatures of S-2A and S-2B. Since this is an externally mounted item who’s I/F is actively maintained at 20°C by heaters, the temperatures are expected to be very similar. Nevertheless, a temperature difference between both satellites of approx. 1.5K is clearly visible. The STRs on S-2A suffer already from degradation of STR baffle thermo-optical properties, leading to slightly higher temperatures w.r.t S-2B.
Figure III-8: Comparison of Star Tracker Temperature between S-2A and S-2B.
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On the Star Trackers of S-2A, an increase of the orbital mean temperature over the lifetime is noticeable. Within one year, the mean temperature rose by about 0.4°C. This is presumed to result primarily from the degradation of thermo-optical properties of the white paint on the Star Tracker baffles. Due to solar incident fluxes, the absorptivity of the white paint increases, leading to increased temperatures of the baffles. This effect is shown in Figure III-9.
Figure III-9: Evolution of Star Tracker Temperature.
V. Conclusion
The satellite thermal design has been analyzed during the different project phases (e.g. PDR, CDR, sensitivity analysis, test prediction, final flight prediction) and verified through thermal balance and vacuum tests to validate the models. The thermal balance performed with the S-2A satellite showed already a good match of the overall thermal model to the as-built status. Nevertheless, some discrepancies were found for example on the Star Tracker Assembly and the Propulsion Plate. The thermal test of the S-2B satellite revealed further needs to improve the thermal model on the already correlated units during the S-2A correlation.
The comparison of flight data during the in orbit commissioning phase with the flight predictions performed with the correlated model show discrepancies in the order of only +/-3K for most units. The comparison of duty cycles shows deviations higher than 10% only for the units which already had higher discrepancies during the thermal balance correlation, namely Star Tracker Assembly and Propulsion Plate. The comparison of flight data reveals a good similarity in temperature, stability and performance of both satellites.
On the other hand, AIRBUS is looking for optimizing both satellite and instrument thermal designs in order to deliver to ESA future flight models (S-2C/D) with improvements for which benefits have to be demonstrated or confirmed. The goal is also to reduce the design margins in order to demonstrate that extending the imaging duration, and generally the S-2 mission, are feasible as well.
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The preliminary in-flight investigations have outlined that:
At system level: conservative requirements between the system and the sub-systems have led to design and/or development constraints that may be relaxed for the future flight models (example: MSI electronic VCU). Comparing TICD and early specifications to real flight measurements will provide lessons learnt on how to improve system and sub-systems exchanges focusing on the Sentinel 2 program and to define more general guidelines for future ESA programs. Comparing S-2A to S-2B is also a good opportunity to identify design improvements and/or robustness.
For both Satellite and Instrument, analyzing in-flight data in more detail will provide a great opportunity to progress significantly on how to design a high-level quality thermal architecture to handle challenging scientific missions.
Acknowledgments The authors want to thank the complete Sentinel-2 Prime and Instrument Team, ESA, ESOC and all suppliers for their support during the design, development and validation process. IABG and ESTEC are acknowledged for their support in the preparation and execution of the TBTV tests.
References
1Cataloglu, A. and Eckert, K., “Launcher Plume Effect on the Satellite”, Proceedings of the 46th International Conference on Environmental Systems, Vienna, Austria, 2016, ICES-2016-114.
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